A telecommunication satellite is placed in mission orbit by the combination of a launcher spacecraft and of its own propulsion means. According to a known technique, diverse service instruments and mission instruments are held against the structure of the satellite in a first configuration, termed the storage configuration. After separation with the launcher spacecraft, these instruments are deployed to an operational configuration allowing their operation. Thus, the implementation of solar generators held against North and South faces of a parallelepipedal structure during a launch phase, and deployed and oriented towards the sun after separation of the launcher craft, is known. It is also known to employ antenna reflectors held against East and West faces of the structure in the storage configuration and deployed so as to allow during the mission the reflection of a beam of waves between a source block fixed to the structure and a zone of coverage of the terrestrial globe.
The increasing of the payload capacity of a satellite within the limits imposed by the nose cone of the launcher spacecraft remains an important issue. Advances in telecommunications services (reduction in the size and power of the user terminals on the ground, geographical reuse of frequencies, related to the sparseness of the spectrum, search for more precise contours formed) involve improvements to the performance of antennas. Employing high focal length antennas, or antenna reflectors of wide diameter, constitutes an avenue of progress. To boost the power of antennas, it is also apposite to increase the dissipative capacity of the satellite so as to optimize the evacuation of heat generated by the mission instruments. More generally, it is sought to increase the area of the surface for rigging facilities on the structure of the satellite, within the limits imposed by the nose cone of the launcher craft.
FIGS. 1a and 1b represent a telecommunication satellite of customary architecture. A satellite 10 is represented in FIG. 1a in an operational configuration allowing the operation of the mission instruments of the satellite in its orbit. The satellite 10 is represented in FIG. 1b in a storage configuration. As represented in FIG. 1b, the satellite can be placed in the interior volume 30 of a nose cone 31 of a launcher spacecraft.
A telecommunication satellite of customary architecture generally comprises a substantially parallelepipedal structure 11 whose orientation is held constant with respect to the earth in the operational configuration. The person skilled in the art uses a reference trihedron tied to the satellite consisting of an axis Z oriented towards the earth, of an axis Y perpendicular to the plane of the orbit, and of an axis X forming with the Y and Z axes a right-handed orthogonal reference frame; the X axis then lying along the direction of the velocity in the particular case of circular orbits.
In a conventional architecture, a face 13 of the structure 11, perpendicular to the Z axis, is commonly called the earth face because of its orientation towards the earth, the opposite face 14 commonly being called the anti-earth face. A face 15 perpendicular to the Y axis and oriented towards the North in the terrestrial magnetic field is called the North face; the opposite face 16 commonly being called the South face. A face 17 perpendicular to the X axis and oriented in the direction of the displacement of the satellite is called the East face; the opposite face 18 commonly being called the West face.
On the North and South faces are customarily fixed solar generators 19 and 20 which ensure the electrical energy supply to the satellite. These latter are motorized so that the surfaces which bear the photovoltaic cells always point towards the sun. The North and South faces also have the particular feature, whatever the position of the satellite in the orbit, of receiving the solar flux with a low or indeed zero incidence. They are therefore used to radiate into space the energy dissipated by the operation of the electrical facilities of the satellite. The other faces receive the solar flux with a high incidence according to the position of the satellite in its orbit. In the storage configuration, the solar generators are folded up and held against the North and South faces so as to limit their bulk and ensure that they are held so as to withstand the dynamic accelerations and the high vibratory stresses of the launch phase.
On the earth face are generally mounted diverse mission instruments, such as for example a gregorian telecommunication antenna 9 such as represented in FIG. 1a. The anti-earth face is generally used to fix the satellite to the launcher. It also generally carries the apogee motor charged with ensuring that the satellite is placed on station as a supplement to the launcher spacecraft.
The East and West faces can be used to rig up antennas. Antennas comprising a radiofrequency source 21 fixed on the structure of the satellite and a deployable reflector 22 such as represented in FIG. 1a are in particular known. In the storage configuration, the antenna reflector is held against an East or West face, it is thereafter deployed by a rotation motion around an axis substantially parallel to the Y axis. In the operational configuration, the reflector 22 is positioned so as to reflect, in an optimal manner, a beam of waves between the radiofrequency source 21 and a targeted terrestrial coverage zone. The radiofrequency sources, associated with reflectors deployed as East or West faces are usually fixed to the structure of the satellite on the East or West faces, or on the edges common to the East or West faces and to the earth face, or else on the earth face in the case of the use of intermediate reflectors, which ensure the reflection of the beam of waves between the source and the deployable reflector.
The current solutions suffer from limits that the present invention seeks to solve. Thus, an antenna reflector held against the structure of the satellite is constrained by the dimensions of the structure of the satellite having to be stored in the nose cone of a launcher craft. Typically the diameter of the rigid reflectors 22 and 23 is generally limited to the dimensions of the faces of the parallelepipedal structure of the satellite. A known alternative solution consists in having unfurlable reflectors consisting of several rigid parts. This type of reflector which generates interference of the beam by the presence on the reflecting surface of uncontrolled reliefs related to the deployment of the various rigid parts of the reflector is in practice little used.
Moreover, the dissipative capacity of a satellite is constrained by the dimensions of the North and South faces. To improve this dissipative capacity, alternative solutions consisting of unfurlable radiators are envisaged. Here again, these alternative solutions exhibit difficulties: complexity and cost of the thermal system, increase in the mass, loss of reliability, limitation of the deployment zones not interfering with the reflectors.